Tubular-shaped launcher for projectiles, in particular for missiles

ABSTRACT

The present invention relates to a tubular-shaped launching device for projectiles, in particular for missiles.

United States Patent 1151 3,653,288

Staufi et a1. [4 Apr. 4, 1972 [54] TUBULAR-SHAPED LAUNCHER FORPROJECTILES, IN PARTICULAR FOR 1 References t d MISSILES UNITED STATESPATENTS [72] Invent: Emile Versailles; Jean Guam" 1 380 35s 6/1921 Cooke..s9/1.7 A

Chatenay-Malabry; Pierre Allard, Fontenay-Sous-Bois, all of France;Johannes Schubert, Putzbrunn, near Munich; Heinz 2,421,522 6/1947 Pope2,598,256 5/1952 Hickman 2 834 255 5/1958 Musser... Tf,Mh-P1'h;EhP, fi fa" 5 i m 2,967,460 1/1961 Mussel... ..s9/1.7 A 2,986,973 6/1961 Waxman...89/1.703 Assigneer Noni-Aviation Societe Nationale De 2,987,965 6/1961Musser ..s9/1.7 A

a or struction Aeronautiques, Paris, France a 9 a [22] Filed: Feb. 25,1965 FOREIGN PATENTS 0R APPLICATIONS [21] pp 435785 460,353 10/1949Canada ..89/1.7A

Primary Examiner-Samuel W. Engle [30] Foreign Application Priority DataAtt0rneyl(ar1W.F10CkS Feb. 26, 1964 France .965 258 [5 ABSTRACT [52][1.8. CI ..89/1.703, 89/1.704 Th ent invention relates to atubular-shaped launching [51 Int. Cl .1 41 de i e for roje tiles inparticular for mi55i1e5 [58] Field ofSearch ..89/1, 1.7 A, 1.7 B, 1.703,

89/ 1 .704 9 Claims, 5 Drawing Figures lullllllllllllgmlllllTUBULAR-SHAPED LAUNCHER FOR PROJECTILES, IN PARTICULAR FOR MISSILES Thepresent invention relates to a tubular-shaped launching device forprojectiles, in particular for missiles.

Devices of this kind are already known which operate in accordance withthe principle of the single chamber or of the double chamber. Accordingto the single chamber conception, the propellant charge is disposedbetween the projectile and a rear closing wall which, as the propellantcharge is consumed, is pushed rearwards, thus ensuring practicallyrecoilless firing.

The single chamber structure also has the advantage of simplicity ofconstruction and of operation.

However, a drawback has manifested itself: the field of use of thissingle chamber design is considerably limited by the minimum pressurewhich is necessary for a stable combustion of the propellant powder. Asa result of the definite limits in the ignition conditions of thepowder, strong accelerations are produced of such a nature that theoperation of a firing device embodying the principle of the singlechamber, for example for the launching of bodies provided with measuringinstruments or equipments which are sensitive to accelerations, is nolonger possible.

An improvement on the single chamber design is achieved by the use of alauncher constructed under the principle of the double chamber andcomprising a high-pressure combustion chamber in which is disposed thepropellant charge, as well as a low-pressure chamber connected to thefirst chamber and comprising an expansion nozzle. This design has theadvantage that all the parameters, such as acceleration of theprojectile, acceleration path, discharge speed, etc. can be alteredwithin wide limits irrespective of the properties of the propellantpowder. The shape and size of the projectile or of the missile launched,as well as the possible existence of control equipment or directionalequipment which are sensitive to accelerations do not limit the field ofuse of this double chamber concept. However, it has been found thatlaunchers operating under the double chamber principle have anunfavorable volume for easy handling of the weapon, either because thediameter of the tube is too large or its length is too long. These twodrawbacks arise from the manner in which the propellant charge isdisposed and from the shape of the expansion nozzle which is usuallydesigned along the lines of a de Laval nozzle.

The object of the present invention is a tubular-shaped launcher of thetype hereinbefore described and exhibiting good possibilities ofadaptation to the highest discharge speeds while having the lowestvolume and weight compatible with the requirement of absence of recoil.

The object of the invention is achieved by means of a highpressurecombustion chamber fixedly secured within the launching tube andprovided with means for receiving solid propergol and at its frontportion with apertures for the passage of combustion gases flowing intoa low-pressure combustion chamber, the wall of the high-pressurecombustion chamber forming with the launching tube a supersonic annularnozzle through which the combustion gases flow.

In contradistinction to known launching devices, the device constructedin accordance with the above principle is characterized not only by asubstantial absence of recoil due to the annular nozzle design but hasthe advantage of a reduced length together with a constant tube diameterover the whole of its length. The reduced length is due to the fact thatin launches of the known type the total length results from the lengthof the nozzle of the combustion chamber and from the length of the spacereserved for the propergol, and to the fact that an annular nozzle ismuch shorter in length than a de Laval nozzle.

To ensure an absence of recoil of the launcher, it is not necessary thatthe forces be in balance at each moment of the combustion. It sufficesthat the sum of the impulses as a whole be zero during the very shortperiod of combustion. In case the thrust through the supersonic nozzlewhich is necessary to balance the quantities of motion, is no longersufficient per se,

according to a unique feature of the invention, the high-pressurechamber has, at its rear portion, openings in the form of reactionpropelling nozzles, which provide an additional thrust. This arrangementhas, furthermore, the advantage of maintaining a small caliber for agiven thrust and pressure within the combustion chamber.

Should the launcher be provided with spin-imparting devices such asguides for example, or any other rotation-imparting means, with a viewto stabilizing the projectile or missile, the angular acceleration ofthe projectile produces a moment of rotation during the firing phase,which acts on the firing installations. On recoil-less firing devices itis however desirable to avoid applying to the tube, during the firingphase, not only forces but also couples. To this end, according to theinvention, there is provided in the annular slot of the supersonic noulea device which produces an antagonistic moment.

In the firing devices according to the invention, making use of thedouble chamber principle involving the production of a thrust in thehigh-pressure chamber and constant sections of flow, the impulses arepractically negligible during the whole of the combustion. However, itis not impossible that under certain conditions, with constant sectionsof flow, the recoil forces are larger than the counter-thrust of thenozzles at the start of combustion, whereas, at the end of combustion,the counter-thrust of the nozzles is preponderant. In this case, theequipment is subjected to axial forces of variable magnitude anddirection. To avoid the production of these forces, ac-

cording to the present invention, the openings provided in the" frontportion of the high-pressure combustion chamber comprise fusible orcombustible gaskets whereby the flow section from the high-pressurecombustion chamber can be increased during the combustion period.

Another solution for avoiding residual impulses consists in providing inthe front portion of the high-pressure combustion chamber openingsproduced by the retarded combustion of certain substances.

In this manner, with either of these solutions, it is possible to adaptthe gas flow to the increase of the tube volume resulting from themotion of the projectile or of the missile within the tube, and toreduce to an appreciable extent the variations of pressure in thelow-pressure chamber, as well as the nonbalanced residual forces.

A form of embodiment of the launching device according to the inventionis shown in the figures of the accompanying drawing. All portions of thedevice which have no direct connection with the invention have not beenshown in detail, for the sake of clarity.

In said figures:

FIG. 1 is a diagrammatical longitudinal section of the extreme end of alauncher;

FIG. 2 is an end view of the launching tube as seen in the direction ofthe arrow A of FIG. 1;

FIG. 3 is a partial view in longitudinal section of the rear part of thehigh-pressure combustion chamber with additional outlet orifices;

FIG. 4 is a view of the forward portion of the combustion chamberprovided with gaskets which produce a variation in the section of theorifices for the passage of the gases;

FIG. 5 is a developed view of the additional gaskets closing theorifices for the passage of the gases in the forward portion of thecombustion chamber.

FIG. 1 illustrates a tube 1 into which a partially shown missile 2 hasbeen introduced. This missile is provided with foldable surfaces 2a forstabilization purposes as is well known in the art. With a view tostabilizing the projectile or part of the projectile by a spin effect,the inner wall of the launching tube comprises guides 1a one of whichonly is shown in the drawing. A piston 3 is disposed at the rear portionof the missile 2, said piston having on its outer periphery rings 3a and3b to ensure a tight seal of the low-pressure chamber in relation to thespace into which the missile is introduced.

Corresponding to the number of guides, the piston is provided on itsperiphery with recesses 30 which is engaged by a block 4 sliding in oneof said guides. In view of the mechanical connection of the piston 3with the guides 1a referred to above, on the one hand, and of thedynamical connection of said piston with the missile, on the other hand,for example through the medium of the shoulder, 3d, as the volume of gasincreases the missile is ejected from the tube with a spin.

In accordance with the invention, the gas (and therefore the thrust) isproduced in a high-pressure chamber 6 fixedly secured to the tube 1 by aring provided with arms 5a; the propellant charge, composed of severaldisks, is mounted within the high-pressure combustion chamber on a tube8 and is secured against axial movements. Rearwardly, the highpressurechamber 6 is hermetically sealed by a disk 9 on which is also securedthe tube 8. At the front of the combustion chamber gas dischargeopenings 6a issue into the intermediary or low-pressure chamber 10arranged rearwardly of the piston 3.

According to the invention, the high-pressure combustion chamber 6 is sodesigned and arranged that its wall 6b, in its central portion, formswith the launching tube 1 a supersonic annular nozzle 11 through whichthe combustion gases flow, whereby the advantages described at thebeginning of the present specification can be obtained.

The firing of the propellant charge 7 is carried out in the illustratedform of embodiment by electrical means (ignition wires 12) which causean igniting device (not described in detail) to detonate. This ignitingdevice is disposed within the tube 8 and ignites the propellant chargethrough apertures (not shown) formed in the tube 8.

In case the caliber must be maintained as low as possible, for apredetermined thrust and chamber pressure, or in case the balancing ofthe quantities of motion obtained by the thrust produced by thesupersonic annular nozzle is not sufficient per se, reactive thrustnozzles 13 can be provided at the rear end of the chamber 6, on the enddisk 9, in addition to the forward orifices 6a for the flow of gases.These reactive thrust nozzles 13 can be designed in the form of de Lavalnozzles or of annular nozzles. An arrangement of this type is shown inFIG. 3.

In order to compensate the moment of rotation resulting from thetorsional couple produced at the start of the missile, and to maintain,as far as possible, the whole launching installation in equilibrium asto moments, the radial fixing arms 5a of the high-pressure combustionchamber are disposed obliquely in relation to the direction of incidentflow (FIG. 2). In this manner, a tangential component is obtained whichis added to the thrust produced by the ejection of the gases discharged,said component giving rise to a couple which is antagonistic to themoment of rotation of the launching tube.

According to FIG. 4, with a view to avoiding axial stresses of variablemagnitudes and directions which are liable to be produced on thelaunching tube in certain cases at the beginning and at the end of thecombustion, the front orifices 6a of the high-pressure combustionchamber are provided with fusible or combustible gaskets 14 which enablethe section of the orifices 6a, which are small at the onset ofcombustion, to be increased during the combustion phase, proportionallyto the increase of the volume in the launching tube during the launchingoperation.

With a view to avoiding these residual forces another form of embodimentis shown in FIG. 5 wherein the front portion of the wall of thecombustion chamber 6b is provided with apertures 6a shown in developedform. As shown in this figure, a plurality of apertures 6a arecompletely closed by gaskets 15 at the start of combustion. During thelaunching operation, a number of additional apertures are formed due tothe combustion or the fusion of said gaskets.

The invention is not limited to the embodiments according to FIGS. 4 or5; on the contrary it is possible to combine both been described andillustrated in an exglanatory manner without any intention to limit thesame, an that alterations 0 detail can be made thereto within the spiritof the invention, without falling outside its scope.

What is claimed is:

l. A tubular launching installation for projectiles comprising alaunching tube having a portion of substantially constant innerdiameter,

a tubular means forming a high-pressure combustion chamber disposedwithin said tube and spaced therefrom and containing a propellantcharge,

said tubular means having forward and rearward closure means andperipheral gas outlet openings adjacent sai forward closure means, a

a supersonic annular nozzle formed by a portion of the outer surface ofsaid tubular means in cooperation with a portion of the inner surface ofsaid launching tube at a region of smallest cross-sectional area betweensaid portions of said surfaces located at the rearward end of saidportion of substantially constant inner diameter of said launching tube,

means for supporting said tubular means in said launching tube locatedin the region of said supersonic annular nozzle,

a piston disposed in said launching tube behind the projectile,

and a low-pressure chamber located behind said piston and in front ofsaid region of smallest cross-sectional area of said supersonic nozzleand concentric to said tubular means.

2. The tubular launching installation of claim 1 further characterizedby said rearward closure means having additional openings in the form ofreaction thrust nozzles.

3. The tubular launching installation of claim 2, further characterizedby guide means on the inner surface of said launching tube for impartinga spin to the projectile,

and said means for supporting said tubular means including gasdeflecting means for producing an antagonistic moment to said spin.

4. The tubular launching installation of claim 1, further characterizedby gaskets in said peripheral gas outlet openings adapted to vanish dueto the heat of combustion during the combustion of the propellantcharge.

5. The tubular launching installation of claim 4, further characterizedby said gaskets being fusible.

6. The tubular launching installation of claim 4, further characterizedby said gaskets being combustible.

7. The tubular launching installation of claim 1, further characterizedby supplemental peripheral gas outlet openings adjacent said forwardclosure means having gaskets adapted to vanish due to the heat ofcombustion during the combustion of the propellant charge.

8. The tubular launching installation of claim 7, further characterizedby said gaskets being fusible.

9. The tubular launching installation of claim 7, further characterizedby said gaskets being combustible.

PatentNo. 3,653,288 Dated April 4, 1972 n (s) EMILE STAUFF, JEANGUILLOT,PIERRE ALLARD,

JOHANNES SCHUBERT, HEINZ TOPFER, AND ERICH PRIER It is certified thaterror appears in the above-identified patent and that said LettersPatent are hereby corrected as shown below:

In column 1, lines 10 and ll, the names of the assignees should read:Nerd-Aviation Societe Nationale de Constructions Aeronautiques Paris,France and Bolkow Gesellschaft mit beschrankter Haftung Munich, WestGermany Signed and sealed this 3rd day of October 1972,

(SEAL) Attest:

EDWARD MQFLETCHER J Attesting Officer ROBERT GOTTSCHALK CommissionerofPatents FORM FO-105O (1069) USCOMM'DC 60376-F59 9 U.S, GOVERNMENTPRINTING OFFICE: 1969 0-366-334 UNITED STATES PATENT OFFICE CERTIFICATE0F CORRECTION Patent N v Dated April 4,

Inventor(s) EMILE STAUFF, JEAN GUILLOT,PIERRE ALLARD,

JOHANNES SCHUBERT, HEINZ TOPFER, AND ERICH PRIER It is certified thaterror appears in the above-identified patent and that said LettersPatent are hereby corrected as shown below:

In column 1, lines 10 and I 11, the names of the assignees should read:Nord-Aviation Societe Nationale de Constrnctions Aeronautiques Paris,France and Bolkow Gesellschaft mit beschrankter Haftung Munich, WestGermany Signed and sealed this 3rd day of October 1972.

(SEAL) Attest:

EDWARD M.FLETCHER,JR.

testing Officer ROBERT GOTTSCHALK Commissioner of Patent:

, FORM PO-IOSO (10- USCOMM-DC 60376-P69 9 U.S. GOVERNMENI' PRINTINGOFFICE: \969 0-356-334

1. A tubular launching installation for projectiles comprising alaunching tube having a portion of substantially constant innerdiameter, a tubular means forming a high-pressure combustion chamberdisposed within said tube and spaced therefrom and containing apropellant charge, said tubular means having forward and rearwardclosure means and peripheral gas outlet openings adjacent said forwardclosure means, a supersonic annular nozzle formed by a portion of theouter surface of said tubular means in cooperation with a portion of theinner surface of said launching tube at a region of smallestcross-sectional area between said portions of said surfaces located atthe rearward end of said portion of substantially constant innerdiameter of said launching tube, means for supporting said tubular meansin said launching tube located in the region of said supersonic annularnozzle, a piston disposed in said launching tube behind the projectile,and a low-pressure chamber located behind said piston and in front ofsaid region of smallest cross-sectional area of said supersonic nozzleand concentric to said tubular means.
 2. The tubular launchinginstallation of claim 1 further characterized by said rearward closuremeans having additional openings in the form of reaction thrust nozzles.3. The tubular launching installation of claim 2, further characterizedby guide means on the inner surface of said launching tube for impartinga spin to the projectile, and said means for supporting said tubularmeans including gas deflecting means for producing an antagonisticmoment to said spin.
 4. The tubular launching installation of claim 1,further characterized by gaskets in said peripheral gas outlet openingsadapted to vanish due to the heat of combustion during the combustion ofthe propellant charge.
 5. The tubular launching installation of claim 4,further characterized by said gaskets being fusible.
 6. The tubularlaunching installation of claim 4, further characterized by said gasketsbeing combustible.
 7. The tubular launching installation of claim 1,further characterized by supplemEntal peripheral gas outlet openingsadjacent said forward closure means having gaskets adapted to vanish dueto the heat of combustion during the combustion of the propellantcharge.
 8. The tubular launching installation of claim 7, furthercharacterized by said gaskets being fusible.
 9. The tubular launchinginstallation of claim 7, further characterized by said gaskets beingcombustible.